Combustion control system



R. L. ROD

COMBUSTION CONTROL SYSTEM Dec. 4, 1962 3 Sheets-Sheet 1 Filed Feb. 2, 1959 F/GIZ INVENTOR. AUfiffif 1. F04 4%!4 Q W ATTOP/Vfy Dec. 4, 1962 R. L. ROD 3,066,482

COMBUSTION CONTROL SYSTEM Filed Feb. 2, 1959 3 SheetsSheer, 2

(ONT/POL 48 F/G' 6' IN VEN TOR.

WQ M AWTO/P/VEV Dec. 4, 1962 R. L. ROD

COMBUSTION CONTROL SYSTEM 3 Sheets-Sheet 3 Filed Feb. 2, 1959 mw U NP. M M WWW N 7 I 0 A f 0 FW W United States Patent Ofifice 3,066,482 Patented Dec. 4, 1962 gas $366,482 CMBUSHON @GNTROL SYSTEM Robert L. Rod, Roslyn Heights, FLY, assignor to Accns: tica Associates, lino, Mineola, N. Y., a corporation of New York Filed Feb. 2, 1959, t er. No. 323,913 15 (Claims. (til. fill-35.6)

This invention relates to systems for controlling the combustion of solid fuels, and more particularly to the control of the burning rate of a solid fuel rocket motor or similar propulsion unit. Such systems are termed propellant utilization systems, or more simply PU systems.

As is well known, a solid fuel rocket motor consists basically of a shaped charge of grain of molded propellant material contained within a metal shell, usually of cylindrical form. it is customary to arrange the propellant along the major axis of the rocket, with a hollow central core provided the full length of the grain. The central core is a bore passing through the entire grain. Such cores may be circular in cross-section, but more often they are star-shaped or have some other complex shape to aid in the control of combustion, as will presently be explained.

The odd-shaped hollow cores are usually known as cruciform cores. They are obtained by casting molten propellant in the cylindrical shell which is fitted temporarily with a solid metal core. When the metal core is removed after the propellant hardens a hollow core is left. During actual operation, combustion takes place along the full length of the core, radially outwardly toward the containing shell, and the combustion gases are exhausted through the exhaust port or nozzle.

The purpose or" the odd-shaped cruciform, when it is used, is to insure that the burning area is changed in some predetermined manner as the combustion proceeds. As is also well known, missile guidance is complicated by non-uniform acceleration during the burning phase of missile flight. Consequently, missile designers usually require acceleration to be nearly constant during the burning phase. Since the mass of the rocket decreases during flight, a regressive time-thrust characteristic is desirable. To accomplish this, the solid propellant motor is constructed with a shaped combustion chamber, the shape being such that the burning surface is greater at the beginning than at the end of the firing. In this man ner, a thrust program is built into the motor. Unfortunately this method precludes any in-flight correction, and the random fluctuations of combustion which are inherent in solid-propellant motors can produce adverse effects (some of which are described below) leading to certain disadvantages, such as decreased missile or rocket range, ballistic errors, and in extreme cases combustion instabilities leading to blow-ups.

Solid propellants are usually formulated by combining an oxidizing agent with a suitable fuel, this mixture generally being in the form of a rubbery material. The term solid when used in this specification and in the appended claims is intended to refer to such rubbery material and all other forms of solid propellants. When conventional solid propellants are initially ignited, by some form of ignitor, the burning gases of combustion seek a release from the combustion area through the orifices or nozzle which is provided. In so doing, the gases, which start out at subsonic velocity, are accelerated to supersonic velocity through the nozzle, and thus pro vide the thrust necessary for propulsion.

For reliable operation it is necessary that the propellant burn uniformly throughout the entire propelling operation. Entrapped bubbles in the propellant, as Well as flaws and cracks in the propellant material itself, tend to cause uneven burning. Frequently thishas catastrophic results. Further, variations in the IIllXll'lg process tend to create unequal distributions of fuel and oxidizer, thereby compounding this difficulty. A prunary disadvantage of these faults is that the containing shell which is usually of steel, and is desirably made as light as possible in order to increase as much as possible the pay load to be propelled by the rocket motor, must be overdesigned to accommodate unexpected pressure peaks. Furthermore, non-uniform or erratic burning complicates the guidance problem, where guided missiles are being propelled, since the guidance equipment cannot be designed to anticipate the erratic changes in propulsion brought about by non-uniform burning. The problem of maintaining a smooth thrust magnitude is further complicated by the rush of hot gases over the grain surface. This not only causes non-uniform burning rates along the entire length of the rocket motor, but also erosive burning, in which chunks of unburned propellant are torn loose and blown out the nozzle, thereby reducing the etficiency of propellant utilization. Presently this is compensated for by forming the combustion chamber in what might otherwise be considered an undesirable shape.

it is an object of my invention to provide a system which controls the burning of solid fuels to burn uniformly. Another object is to provide such a system which assures that the burning pressure is maintained about a preset value, and according to a prescribed thrust program, if desired. It is a further object of my invention to provide a system which prevents unexpected pressure peaks during the combustion of solid fuels, thereby eliminating the need to overdesign the containing shells, so that rockets and missiles may be built of lighter materials and with increased range capabilities, and with pre dictable guidance characteristics. It is a further object of my invention to increase the rate of heat transfer between combustion gases and unburned fuel, and thereby increase the thrust available from a given charge of fuel. A further object of my invention is to eliminate the defects caused by non-uniformly cast propellants and by propellants which have flaws in them including air bubbles and cracks, and to make possible the use of all such propellants to provide the optimum performance when they are burned. A general object of my invention is to provide a combustion control system for solid fuel rocket motors capable of maintaining a predetermined thrust program without the use of special core or combustion chamber shapes, and one which will provide more efiicient propellant utilization, and which will permit instantaneous automatic in-flight thrust control to be maintained.

The general nature of the invention is a system of propellant utilization which overcomes the aforementioned disadvantages and provides an added measure of combustion efficiency through the use of elastic wave or vibratory energy. According to the invention a source of elastic wave or vibratory energy is arranged or adapted to supply such energy to the region in a solid fuel propellant container in which combustion is intended to take place, together with means to control said source so that vibratory energy actually controls the combustion in a predetermined manner. The invention is .applicable to solid fuel propellants which are cast with a circular or other shaped core, as well as to propellants which have no hollow interior core, but rather are of the end burning type. The invention makes use of all forms of sources of vibratory or elastic wave energy, including gas driven devices such as sirens and whistles as well as electromechanical transducers, and where hereinafter the term transducer appears, whether in the specification or in the appended claims, that term is intended to inaoeaasz clude all such forms. The invention contemplates that the transducer may be driven by an electrical source, by compressed gases including air, or even by burning gas from the combustion of the fuel being controlled itself. as well as by combinations of any or all of these means. More specifically, the invention in one embodiment includes a sensing device, which for example may be placed in the exhaust nozzle of a rocket motor, which device measures a characteristic of the burning gases, for example the output velocity, the temperature or some other characteristic, and provides a signal to a control device which regulates the amount of elastic wave energy which is supplied to the burning fuel. In another en bodiment of the invention, the sensing device is respo sive to the acceleration of the vehicle being propelled by the rocket motor, and the control signal is a function of such acceleration. The elastic wave energy employed may be of the continuous wave type, including so-called white noise, or an output of a high harmonic content, as well as single frequency content elastic Wave energy; further, the latter includes the employment of discreet frequency wave energy, which may be of extremely high intensity, together with means to vary the frequency so that standing wave peaks having certain advantages to be described below may be employed, and may be shifted as desired throughout the entire volume of combustion so that certain desirable features provided by such standing wave peaks may be uniformly distributed throughout the combustion region.

Other and further objects and features of the invention will become apparent from the following description of certain embodiments thereof. This description refers to the accompanying drawing wherein:

FIG. 1 illustrates a system according to the invention;

FIG. 2 is a modification of FIG. 1;

FIG. 3 is a diagram illustrating the nature of burning of a solid fuel;

FIG. 4 is another embodiment of the invention, employing a gas driven elastic wave generator;

FIG. 5 is a variation of FIG. 4;

FIG. 6 is still another variation of the device of the system shown in FIG. 4;

FIG. 7 illustrates an embodiment of the invention employing combustion gases to drive a wave generator; and

FIG. 8 is a diagram relating to gas port shapes.

Referring now to FIG. 1 a container ltl, here shown as being of cylindrical form, has within it a solid propellant 11 molded into the container with a central circular core 12 provided for the combustion area. A cir cular core is shown for simplicity, although it will be realized that any odd-shaped core or even a cruciform core is equally suitable. The container is fitted with a nozzle 14 at the rear, which accelerates the burning gase which are exhausted during the combustion process. in the forward part of the container there is located a solid plate 15 which contains the combustion and which for the purposes of this invention is made of a material which is capable not only of containing the combustion but is also substantially transparent to elastic wave or vibratory energy. Any stainless steel of the 360 series is suitable for use as this plate 15, as is any other wavetransparent material which is capable of withstanding the high temperatures of solid-fuel combustion for the period of three to four minutes of burning time usually encountered in rocket motors.

Mounted on the exterior of the wall 15 is a source of elastic wave energy, comprising a generator of elastic wave energy 17 and a born 18, through which the elastic wave energy generated by the generator 17 is applied through the wall 15 to the core 12.

A source of driving energy for the generator 17 is shown at 20. This source feeds through a control unit 21 to the generator 17. A part of the output pressure, velocity, acceleration, temperature or other characteristic existing in the exhaust gases is measured by a sensing device 22 which is located in the nozzle 14, and the signal generated by the sensing device is applied to the control unit The sensing device 22 and control unit comprise a feedback signal loop, and these are employed to control the output of the generator 1'7 in applica ions where the thrust (or burning pressure) is to be controlled in some predetermined manner with respect to time. inasmuch as feedback techniques are well known, no specific feedback elements are herein described or illustrated, but it will be understood that any suitable embodiments of the diagrammatically illu"- trated source 24), control unit ill and sensing device 22 may be employed. The sensing device 22 should be made of materials which are able to withstand the con- "ions present in the nozzle 14 during the burning phase.

however, the characteristic being sensed is acceleration, the accelerometer which is usually incorporated in a missile or rocket for other purposes can be used to supply the signal to the control unit and the sensing device 22 can be eliminated as superfluous. The illustration of a sensing device in FIG. 1 is intended to exemplify all such possibilities.

As is illustrated in FIG. 2, the feedback loop may be omitted if desired. This will be done for example in applications where only the advantages resulting from an improvement in heat transfer are desired, in which case the source 2%? will drive the generator 17 directly at a preset intensity value controlled by a control unit 24 which for example can be manually preset. As will be appreciated, in the more general case, there will be utility for and therefore it will be desirable to employ the feedback system which is shown in FIG. 1.

Two discreet forms of elastic wave energy are feasible for use in accordance with this invention. The first is so-called white noise, or an elastic wave energy output of high harmonic content. in this case elastic wave energy from the generator 17 generally promotes heat transfer in the core 12 between the inner surface of the propellant and the gases (not shown) in the core. This mechanism is believed to be as follows.

Referring to FIG. 3, present combustion theory depicts the propellant burning process as occurring in at least three separate phases progressing outward from the propellant surface. FIG. 3 illustrates a small portion of the propellant 11 above which (in the figure) there are shown, progressing outwardly from the surface 31 of the propellant, first a foam zone 32, followed by a fizz zone 33, which is followed by a preparation zone 34, and finally by the flame zone 35. The direction of exhaust gases is indicated by an arrow 36, which is intended to show the direction toward the nozzle 14 in MG. 1. The direction of heat transfer, indicated by an arrow 37, is from the flame zone 35 to the propellant 11 through the preparation zone, the fizz zone and the foam zone. The burning rate of the propellant is proportional to its surface temperature Ts, which in turn is a function of the subsurface foam reaction (which is exothermic) and 01. the heat transfer from the hot central flame (that is, from the flame zone 35) to the relatively cooler surface 31 of the propellant ll, by thermal conduction from the gas phase. The greater the rate of such heat transfer, the greater will be the burning rate. That is, an increase in heat transfer from the gas phase to the surface 31 will lead to an increased burning rate, i.e., more emcient combustion, and increased thrust. The foregoing information in this paragraph will be foundin USED 6374, volume 1Unsolved Problems in Solid Propellant Combustion, by Richard D. Geckler, in the Fifth Symposium (international) on Combustion.

The gas-phase fizz and flame reactions affect the burning rate only as they affect Ts, through such thermal conduction. At low pressures, little heat returns from the gas phase, and the surface temperature approaches a limiting value, Ta, determined essentially by the heat evolved in the subsurface foam reaction. Thus, at low pressures the burning rate will tend to be constant at the value corresponding to Ta. The fizz and flame reactions together establish a temperature gradient in the gas phase. As the pressure increases, the return of heat from the hotter zones to the cooler zones (i.e. from the flame Zone 35 to the foam zone 32 and the surface 33.) will be greater, causing Ts to increase with increasing pressure. The burning rate will increase with Ts. The temperature gradient of the gas phase may be thought of as moving in against the burning surface as the pressure increases; the steeper the temperature gradient. the higher the pressure index. Thus, a lower pressure index may be achieved by flattening out the temperature gradient, either by lowering the flame temperature itself, or by slowing down the gas-phase reactions so that the same fiame temperature is reached over a longer distance.

The thrust of a solid-fueled rocket motor is a function of the fuel density, the specific impulse of the fuel, the shape of the charge, and the burning rate. The burning rate alone of these parameters is dependent upon the dynamic condition of the burning propellant, so that, for any particular set of operating conditions, thrust is directly proportional to burning rate. As seen above, the burning rate is a function of surface temperature (cg, at surface 31 in FIG. 3), which in turn is proportional to the rate of heat transfer from the hot flame zone to the relatively cooler surface 31.

I have determined that the application of elastic wave energy at an intensity level in excess of approximately 118 decibels to a solid propellant rocket motor will result in a marked increase in the temperature Ts at the surface 31, and therefore in the thrust of the motor. At an intensity level of approximately 170 db the heat transfer coefificient will be approximately double that found at 118 db. These determinations were made on the assumptions that (a) The surface (31) of the burning propellant (Iii) is in the liquid phase, and

(b) The flame (35) remains laminar regardless of pressure increases.

These conditions have been found by observation to exist at the surface of a burnim grain (such as the surface 31. of propellant H in P16. 3). it is known, on the other hand, that the quantity of liquid evaporated per unit time from a liquid surface is mainly proportional to the turbulence of the gas above the surface and to the surface area of the liquid, and is inversely proportional to the gas pressure surrounding the liquid boundary. The phenomenon of accelerated drying through increased gas turbulence due to the application of gas-borne sonic energy is well known. Since in a solid propellant rocket motor the region close to the surface (it) is relatively free from turbulence, the application of elastic wave energy to the motor will create turbulence in that region, and will increase the rate of evaporation, which in turn will directly increase the burning rate and the thrust available from a given grain. Elastic wave energy at the levels mentioned above will have a marked effect. For example, a power level of 160 db, at a frequency of l l(C./SBC. in air, corresponds to a velocity amplitude of the gas molecules above the liquid of 628 cm./ sec.

The foregoing efiects occur with composite as well as double-based propellants. Composite propellants are non-homogeneous and are made up of alternate layers of oxidizer and fuel. The combustion process involves mixing oxidizer and gaseous fuel in the reaction zone so that the burning rate is also a function of the mixing rate. The turbulence created by elastic wave energy increases the mixing rate, and results for this additional reason in an increased burning rate.

The second form of elastic wave energy which may be employed is that in which the generator 17 is constrained to emit a discreet frequency output, preferably of extremely high intensity. For the purposes of this discussion such an intensity may for example be db or more referred to the accepted zero db level of 0.006 watt, but it is, of course, in no way limited to this value. In the case of an axially symmetric grain ll of the type shown in FIG. 1, the core 12, which is gas-filled during combustion, forms a cavity having a natural resonant frequency dependent upon its length, the composition and temperature within the core, and in this case upon the velocity of the combustion gases contained within the core. By proper selection of the output frequency of the source 2%} it is possible to create an acoustic standing wave longitudinally in the core, and this will set up a series of acoustic or elastic wave pressure nodes and anti-nodes along the cores axis (not shown). In the simplest case where the frequency is selected to develop a half wave resonance along the core, one such node and one such anti-node will be established. Operating on a second harmonic basis, for example, one will find a typical pressure distribution as is shown the pressure versus length curve 39 which is drawn just above FIG. 1. Once such a standing wave system is set up, there will be achieved maxima of combustion at the points of maximum pressure, or anti-nodes, and reduced burning efficiency on either side thereof toward the adjacent nodes, this reduction being a sinusoidal function of distance away from the anti-nodes.

This standing wave system can readily be employed to correct the form of combustion instability known as re onance burning, which is sometimes encountered in rocl et motors. Resonance burning takes form of high frequency sinusoidal pressure oscillations which spiral around in the combustion chamber. it can cause sudden rises in chamber pressure. Resultant pressure peaks may be accompanied by break-up of the propellant grains, which in extreme cases can cause the motor to explode. This form of combustion instability can be controlled by applying the elastic standing-wave system to induce a thrust fluctuation in phase opposition to the thrust fluctuation induced by the resonance burning. Phase-sensing circuits suitable for this purpose are well known.

In the regions of minimum pressure, or nodes, the burning might be assumed to be least efficient, but there is a second order effect which tends to alter this condi' ion toward at more favorable combustion rate. it is well known that acoustic energy, applied to a resonant cavity, will bring about an agglomeration of aerosols directed through the cavity, in such a manner that tie particles borne in an aerosol tend to collect at the nodes of acoustic or elastic wave pressure. Particles of certain size are excited by the vibratory energy and move violently and collide with an agglomerate with their neighbors. See for exan.- ple US. Patent No. 2,535,700 to Seavey. Such particles then tend to gather at the nodes and fall out due to the action of gravity. in the present application, .rowcver, unburned propellant particles in the submicron to the ten micron range will tend to gather in the regions of the nodes rather than being blown out by the combustion gases. This is the particle range most readily excited by practical acoustic or elastic wave sirens or whistles. This agglomerating action will tend to minimize the effects of erosive burning and to improve combustion by retaining unburned fuel particles in the core instead of permitting them to be blown out, thereby providing that they will be burned and their energy applied to the propulsion or power process.

Should a standing wave system of nodes and anti-nodes few in number be set up in the core 12 the burning would be accelerated at fixed positions along the grain 11 but not as much accelerated at other positions. By modulating the frequency of the source 2d, that is, by constituting this source a swept frequency source, circuits for doing which are well known in the art, the standing Wave system, which is illustrated by way of example by curve 39, can

c eeses be made to shift position repetitiously along the axis of the core 12, thus moving the nodal and anti-nodal positions rapidly about until the entire core has been exposed to points of maximum pressure and therefore has been exposed to maximum combustion. This is particularly feasible when the system is designed so that the source 20 oscillates at a high order of harmonic resonance for the core 12. For example, if the core is resonant longitudinally at the tenth harmonic of the fundamental fre quency, there will be initially a standing wave system of twenty nodes and twenty anti-nodes, so that a shift in frequeny will not disable the resonant qualities of the core, and furthermore, there will be a large number of anti-nodes throughout the core thus requiring smaller shifts in frequency to accomplish the desired results. Still further, if the frequency modulation is applied at a high rate, the vastly-increased burning then becomes essentially uniform everywhere along the core over a given period of time.

The application of elastic wave en rgy in either of t e aforementioned forms according to the invention will tend to correct another disadvantage resulting from erosive burning, namely, that of thinning the burning zone at one end of the grain relative to the other. This comes about as the burning gas is accelerated toward the nozzle, resulting in the grain burning in a tapered shape. As the erosive burning diminishes the thickness of the burning zone toward the exhaust end the elastic wave energy will exert a greater etfect at the relatively less turbulent head end. This will have a smoothing efiect, and will tend to promote more uniform burning and to decrease such a taper.

it is thus seen that my system of controlling the combustion of solid fuels by means of elastic wave or vibratory energy is useful with energy of many forms, including so-called White noise and discreet frequencies, and becomes particularly efiicient when combined with a feedback loop as shown for example in H6. 1 to control the burning intensity about a predetermined level. Then, should the propellant burn more rapidly than desired, the elastic Wave energy level is automatically re duced to compensate for this er'fect. On the other hand, if the combustion is below normal, an increase in elastic wave energy intensity can be brought about to restore the combustion performance desired. To accomplish this, it is necessary to operate the transducer at some intermediate level, for example 140 db, and then cause amplitude modulation about this normal level correspond ing to the fluctuations in thrust.

FIG. 4 illustrates a siren or whistle it} driven by a com pressed gas (not shown) by means of a pump 41 through a variable resistance element or pressure regulating ele ment 42. The siren or whistle 4b is exemplary of a pure 1y mechanical or hydrodynamic vibrations generator or transducer and it is coupled to the born 18, to illustrate a manner in which it may be coupled to the core 12. in FIG. 1 for example. An elastic Wave generator which combines the features of the siren or Whistle 4t and the horn 18 is illustrated and described in US. Patent No. 2,755,767 issued July 24, 1956, to Robert Levavasseur. The variable resistance 42 is used to adjust the intensity of vibrations produced by the transducer ti. Gas is brought to the pump 41 through an input pipe 4-3 from any desirable source. This source might include the combustion gases from the rocket motor itself by suitable bleeding arrangements, as shown in FIG. 7.

Referring now to PEG. 5, there is shown here, in part, a system for applying a servomechanism according to FIG. 1 to the gas driven generator system of FIG. 4. Only the variable resistance 42 of FIG. 4 is shown. To this resistance there is connected a relay shown generally at 43 which has a coil 44 and an armature 45 for controll ing the variable resistance 52 in accordance with an elec trical signal. A sensing device 224} (which may be the sensing device 22 of FIG. 1) brin s its signal to a control circuit 46 of the kind which is adapted to actuate the relay when the signal brought to it from the sensing device is above a preset level, and to drop the relay out when the signal from the sensing device drops in intensity below such preset level. Circuits of the type which will satisfy the requirements of the control circuit 46 are Well known and need not be illustrated. As indicated above, the sensing device 220 may be the accelerometer of the vehicle being propelled by the rocket motor of which the thrust is being controlled. By the system illustrated in 5, the intensity of combustion of the fuel 11 can be controlled in an on-otf fashion about a preset level.

Referring to FIG. 6 the pump 41 of FIG. 4 is illustrated as driven by an electric motor 47 through a shaft 48. A speed control circuit 49 is supplied an input signal by the sensing device 220 and thereby is adjusted to control the speed of the motor 47. The motor 47 in turn controls the speed of the pump 4-1 to effect continuously variable pressure of gas applied to the transducer 4-6 (FIG. 4-).

in H8. 7, the rocket motor of FIG. 1 is shown in part; like parts of these two figures bear like reference characters. A generator 56 of elastic wave energy, which for the purpose of illustration may be the one shown in the aforementioned Patent No. 2,755,767, is arranged to supply such energy to the combustion chamber 12 through the sound transparent end wall 15. Gas to drive the generator is taken from the combustion chamber over a pipe 5'7 through a pressure regulator 61 and a valve modulator 62 to the input pipe 58 of the generator (corresponding to input pipe 12 of said Patent No. 2,755,767). Gas exhausted from the generator 56 is returned to the combustion chamber 12 over an exhaust pipe 59, which connects to the chamber at a point further removed from the head end of the motor than the driving input pipe 57. The sensing device 2243 (as in FIG. 5 or 6) supplies its signal to a comparator 63, which compares it with a reference signal from a reference signal source 64, and supplies an error signal (e.g., difference between input signals to comparator) to the modulator 62 over an error signal line 65. The valve modulattor may be similar to the variable resistance 42 shown in FIGS. 4 and 5.

The embodiment of FIG. 7, like that of FIG. 1, functions as a closed-loop servo system. Any tendency for the thrust to fluctuate is sensed by the sensing device 224 through corresponding variation in the parameter it is adapted to supervise, and is immediately counteracted by appropriate increase or decrease in the intensity of elastic wave energy being supplied to the combustion chamber if desired, the reference signal can be made to vary in accordance with the desired thrust-time relationship. In this case, corrections will be applied to maintain the established relationship.

As was mentioned above, in many solid-fueled rockets, it is desirable to maintain a fairly constant chamber pressure throughout the period of powered flight. To facilitate this, it is common practice to shape the gas port (chamber 12 in FIG. 1) in such a manner that the burning surface remains constant throughout burning. This is usually accomplished by making the initial gas port cross-sectional shape in the form of a star, cross or other complicated geometry. FIG. 8A shows such a shape of gas port 122, in the cross-sectional view of a grain 111. This method, while effective in maintaining a constant chamber pressure, as illustrated by the pressure vs. time graph 113 in FIG. SA, has inherent disadvantages, some of which are:

(a) A complicated form is difficult to manufacture and inspect;

(11) Approximately 5% of the propellant remains unburned;

(c) The ratio of propellant weight to total motor weight is reduced;

(d) The burning time is reduced;

(e) The sharp pointed parts of the grain are easily torn loose by the exhaust gases; and

(f) The sharp pointed parts are subject to fracture from resonance burning.

A round gas port, as shown at 212 in the grain 211 in FIG. SB, eliminates these disadvantages, but the pressure vs. time curve 213 of such a port is not constant, as is also illustrated in FIG. 8B. As burning progresses, the burning surface increases, and with it the burning pressure, and the thrust. This change in burning pressure can be corrected by controlling the burning rate according to a prescribed time program, as is indicated in FIG. 8C. The burning rate is caused to decrease with time, along the dotted-line slope 214, as the burning surface around the port 212 increases, the slope being such that the pressure vs. time curve 215 remains constant.

As has been seen, the burning rate is controllable by means of elastic Wave energy, and control of the burning rate allows control of the chamber pressure, By choosing an appropriate time-variation characteristic for the reference signal in FlG. 7, a constant pressure vs. time, and hence a constant thrust-time relationship can be established and maintained for a rocket propulsion motor having a simple circular gas port. To accomplish this, the elastic wave generator is programmed to be operated from a high power level at the start of flight to a lower power level as the burning surface area increases, according to the curve 214.

Many other variations of the system according to my invention are feasible. Other types of transducers can be employed. The frequency and intensity of the energy supplied by the transducer 17 in FIG. 1, or 40 in FIG. 4, or 56 in FIG. 7, or any other transducer which is used, are related to the dimensions of the particular grain being considered, and its burning pressure. The degree of correction required to maintain a predetermined burning pressure will be important to establish the median intensity level for the transducer, and its amplitude modulation characteristics. The rate of the frequency modulation, in cases where a swept frequency source of energy is used, which is needed to smooth out pressure fluctuations in cases of a single frequency source and transducer, is again a function of the particular propellant mix used. Generally, varying the frequency of a fixed frequency driving source at for example a 4 cycle per second rate, will be more than adequate for these purposes. The choice of transducer will depend upon the application and is not restricted to any particular type. It is not necessary to use a transducer capable of withstanding the high pressures involved, although a transducer capable of withstanding the high temperatures which will exist for a short period of time in the vicinity of solid fuel combustion is preferred. The wave energy is coupled into the reaction chamber or core 12 through a suitable acoustically transparent window in the wall 15, which is itself capable of containing the reaction pressures. Such a window may be made of stainless steel of the 360 series, as is mentioned above, for example.

My invention is applicable also to the entirely different group of solid propellant motors using end-burning grains. Referring to FIG. 1 as an example, in such a case the housing to is completely filled with propellant 11 and the burning takes place perpendicular to the long axis of the rocket or motor, starting at the nozzle end 14 and Working toward the forward end (at end plate 15). Elastic wave energy can then be supplied to the combustion region and controlled in any of the manners herein taught, including arrangements for programming wave intensity as combustion conditions are altered with respect to time.

Systems according to the invention will also control the form of combustion instability known as chuifing. When the pressure in a rocket motor falls below a certain critical value, the chamber pressure may fall suddenly to atmospheric pressure at which time the fuel charge apparently ceases to burn. After a delay of the order of a fraction of a second to a few or many seconds the charge reignites and burns normally for a period. Cyclic repetition of this process is called chuffing.

The final stages of the gas phase reaction occur some distance from the burning surface, and are relatively sluggish at low temperatures. Interruption of this reaction drops the temperature and amount of gaseous product suddenly, and causes a correspondingly sudden drop in the chamber pressure, which in turn leads to a decrease in energy transfer to the burning surfacethe surface temperature Ts falls and burning ceases. This is followed by reignition due to local heat in the motor walls and in the fuel grain itself, which promotes exothermic reactions in the grain. A combustible ga mixture is released which ignites spontaneously when a critical temperature is reached. The hotter surface layers burn rapidly, raising the chamber pressure to a level high enoughfor the flame reaction to take place. But as soon as these hotter layers are burned away, the burning rate decreases, the chamber pressure falls, and the cycle repeats itself. Systems according to the invention inhibit chufiing, and extend the useful burning time of the solid fuel. Operating the transducer about some intermediate energy level causes an increase in the average energy returned to the fuel surface. Since the energy returned to the surface is the controlling factor for chumng, operation with elastic wave energy according to the invention results in a lowering of the critical temperature and pressure.

The embodiments of the invention herein described are exemplary only. Those skilled in the art will recognize other embodiments and applications of the invention. No attempt has been made herein to describe all possible embodiments, but only to illustrate the best way now known to practice it. It is not intended that the claims which follow be limited to the described embodiments.

What I claim is:

1. System for controlling the combustion of a solid fuel comprising a mass of solid fuel having a surface a jacent which combustion is intended to take place, a source of vibratory energy adapted to supply such energy to the region adjacent said surface in which said combustion is intended to take place, means to drive said source, and means to control said driving means.

2. System for controlling the combustion of a solid fuel comprising a mass of solid fuel having a surface adjacent which combustion is intended to take place, a source of vibratory energy adapted to supply such energy to the region adjacent said surface in which said combustion is intended to take place, means to drive said source, and control means responsive to the results of said combustion to control said driving means.

3. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port, .a solid fuel mass in said container, a bore in said fuel mass opening to said port, a source of vibratory energy adapted to supply such energy to the interior of said bore, and means to control said source.

4. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port, a solid fuel mass in said container, a bore in said fuel mass opening to said port, a source of vibratory energy adapted to supply such energy to the interior of said bore, means to drive said source, and means responsive to the results of said combustion to control said driving means.

5. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port and a wall which is able to contain combustion and is substantially transparent to vibratory energy, a solid fuel mass in said container, a bore through said fuel mass opening at one end to said port and at the other end 11 to said wall, a source of vibratory energy located outside said container adjacent said Wall in a position to supply such energy to the interior of said bore through said wall, and means to control said source.

6. System for controlling the combustion of a solid fuel comprising a fuel container for a mass of solid fuel having an exhaust port, an energy transducer adapted to supply elastic wave energy to the region in said container where combustion is intended to take place, a device located in said exhaust part responsive to a condition produced by said combustion, a source of driving energy for said transducer, and means under control of said condition responsive device to control said source.

7. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port, a solid fuel mass in said container, a bore in said fuel mass opening to said port, a transducer adapted to supply elastic wave energy to the interior of said bore, a device responsive to a condition of burning exhaust gases located in said exhaust port, a source of driving electrical energy for said transducer, and means under control of said condition responsive device to control said energy.

8. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port and a wall which is able to contain combustion and is substantially transparent to elastic Wave energy, a solid fuel mass in said container, a bore through said fuel mass opening at one end to said port and at the other end to said wall, a transducer located outside said container adjacent said Wall in a position .to supply elastic wave energy to the interior of said bore through said wall, and means to control said transducer.

9. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port and a Wall which is able to contain combustion and is substantially transparent to elastic wave energy, a solid fuel mass in said container, a bore through said fuel mass opening at one end to said port and at the other end to said well, a transducer located outside said container adjacent said Wall in a position to supply elastic Wave energy to the interior of said bore through said wall, means to supply energy to drive said transducer from the combustion being controlled, and means under control of a condition produced by said combustion to control said energy.

10. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port, a solid fuel mass in said container, a bore in said fuel mass opening to said port, a gas-driven generator of elas tic wave energy adapted to supply such energy to the interior of said bore, a source of gas under pressure adapted to drive said generator, and means under control of a condition produced by said combustion to control said pressure.

11. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port, a solid fuel mass in said container, a bore in said fuel mass opening to said port, a gas-driven generator of elastic wave energy adapted to supply such energy to the interior of said bore, a gas connection from said bore to said generator providing a source of gas under pres- 12 sure adapted to drive said generator, and means under control of a condition produced by said combustion to control said pressure.

12. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port and a Wall which is able to contain combustion and is substantially transparent to elastic wave energy, a solid fuel mass in said container, a bore in said fuel mass opening at one end to said port and at the other end to said Wall, a gas-driven generator of elastic wave energy located outside said container adjacent said Wall in a position to supply elastic Wave energy to the interior of said bore through said wall, and means to control the output of said generator.

13. System for controlling the combustion of a solid fuel comprising a fuel container having an exhaust port and a wall which is able to contain combustion and is substantially transparent to elastic wave energy, a solid fuel mass in said container, a bore in said fuel mass opening at one end to said port and at the other end to said wall, a gas-driven generator of elastic Wave energy located outside said container adjacent said Wall in a position to supply elastic wave energy to the interior of said bore through said wall, a source of gas under pressure adapted to drive said generator, and means to control said pressure about a median level.

14. System for controlling the combustion of a solid fuel comprising a source of vibratory energy adapted to supply such energy to the region adjacent said fuel in which combustion is intended to take place, means to control said source to provide vibratory energy to said region, means responsive to a condition resulting from said combustion to provide a first signal to said control means, and means to provide a prescribed reference signal to said control means, said control means providing a control signal which is a composite of both said first and said reference signals.

15. System for controlling the combustion of a solid fuel comprising a source of vibratory energy adapted to supply Such energy to the region adjacent said fuel in which combustion is intended to take place, means to control said source to provide vibratory energy to said region, means responsive to a condition resulting from said combustion to provide a first signal to said control means, and means to provide to said control means a second signal varying with respect to time in accordance with a prescribed program, said control means providing a control signal which is a composite of both said first and second signals.

References Cited in the file of this patent UNITED STATES PATENTS 2,549,464 Hartley Apr. 17, 1951 2,573,536 Bodine Oct. 30, 1951 2,767,783 Rowell et al. Oct. 23, 1956 FOREIGN PATENTS 1,136,328 France Dec, 29, 1956 1,155,524 France Dec. 2, 1957 214,616 Great Britain Jan. 29, 1925 763,014 Great Britain Dec. 5, 1956 

